Blade having porous, abradable element

ABSTRACT

A blade includes an airfoil having a base and a free, tip end. The tip end includes at least one porous, abradable element.

BACKGROUND

This disclosure relates to an airfoil, such as an airfoil for a gasturbine engine.

Turbine, fan and compressor airfoil structures are typicallymanufactured using die casting techniques. For example, the airfoil iscast within a mold that defines an exterior airfoil surface. A corestructure may be used within the mold to form impingement holes, coolingpassages, ribs or other structures in the airfoil. The die castingtechnique inherently limits the geometry, size, wall thickness andlocation of these structures. Thus, the design of a traditional airfoilis limited to structures that can be manufactured using the die castingtechnique, which in turn may limit the performance of the airfoil.

SUMMARY

A blade according to an exemplary aspect of the present disclosureincludes an airfoil that has a base and a free, tip end. The tip endincludes at least one porous, abradable element.

In a further non-limiting embodiment of the above example, the tip endincludes a platform and at least one porous, abradable element is on theplatform.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element includes a first porous,abradable element and a second porous, abradable element that definesides of a channel, with the platform defining a bottom of the channeland the channel having an open top opposite the bottom.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element includes a honeycomb.

In a further non-limiting embodiment of any of the foregoing examples,the honeycomb has cell walls that are tapered.

In a further non-limiting embodiment of any of the foregoing examples,at least one porous, abradable element includes a random array of pores.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element includes geometric pores.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element includes a first porous,abradable element and a second porous, abradable element spaced apartfrom the first porous, abradable element.

In a further non-limiting embodiment of any of the foregoing examples,the first porous, abradable element and the second porous, abradableelement are elongated along the same direction and offset from eachother along the direction.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element has a first composition andthe airfoil has a second composition that is equivalent to the firstcomposition.

In a further non-limiting embodiment of any of the foregoing examples,the at least one porous, abradable element has a first composition andthe airfoil has a second composition that is different than the firstcomposition.

A turbine engine according to an exemplary aspect of the presentdisclosure includes, optionally a fan, a compressor section, a combustorin fluid communication with the compressor section, and a turbinesection in fluid communication with the combustor. The turbine sectionis coupled to drive the compressor section and the fan. At least one ofthe fan, the compressor section and the turbine section includes aplurality of blades. Each of the plurality of blades has an airfoil thatincludes a base and a free, tip end. The tip end includes at least oneporous, abradable element.

A further non-limiting embodiment of any of the foregoing examplesincludes a static shroud extending circumferentially around theplurality of blades adjacent the porous, abradable elements, and thestatic shroud includes at least one edge extending circumferentially andradially inwardly toward the porous, abradable elements.

In a further non-limiting embodiment of any of the foregoing examples,at least one edge is circumferentially continuous over multiple ones ofthe plurality of blades.

A method for processing a blade according to an exemplary aspect of thepresent disclosure includes depositing multiple layers of a powderedmetal onto one another, joining the layers to one another with referenceto data relating to a particular cross-section of a blade, and producingthe blade with an airfoil including a base and a free, tip end. The tipend includes at least one porous, abradable element.

A method for processing a blade according to an exemplary aspect of thepresent disclosure includes providing a blade having an airfoilincluding a base and a free, tip end, and forming at the tip end atleast one porous, abradable element.

A further non-limiting embodiment of any of the foregoing examplesincludes prior to the forming, removing a prior porous, abradableelement from the tip end.

In a further non-limiting embodiment of any of the foregoing examples,the removing includes machining the tip end to form a flat surface.

In a further non-limiting embodiment of any of the foregoing examples,the forming includes forming the at least one porous, abradable elementon the flat surface.

In a further non-limiting embodiment of any of the foregoing examples,the forming includes depositing multiple layers of a powdered metal onto one another and joining the layers to one another with reference todata relating to a particular cross-section of the at least one porous,abradable element.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure willbecome apparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

FIG. 1 shows an example gas turbine engine.

FIG. 2 shows a perspective view of an airfoil.

FIG. 3A shows a honeycomb of a porous, abradable element.

FIG. 3B shows a cross-section through a honeycomb that has cell wallsthat are tapered.

FIG. 3C shows a porous, abradable element with a random array of pores.

FIG. 4 shows a plurality of airfoil blades and a static shroudcircumferentially extending around the airfoil blades.

FIG. 5 shows a cutaway view of a static shroud and airfoil blades havingporous, abradable elements.

FIG. 6 shows a method of processing an airfoil.

FIG. 7 shows a perspective view of a tip end of an airfoil having anopening there through.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a first spool 30 and a second spool 32mounted for rotation about an engine central axis A relative to anengine static structure 36 via several bearing systems 38. It should beunderstood that various bearing systems 38 at various locations mayalternatively or additionally be provided.

The first spool 30 generally includes a first shaft 40 thatinterconnects a fan 42, a first compressor 44 and a first turbine 46.The first shaft 40 may be connected to the fan 42 through a gearassembly of a fan drive gear system 48 to drive the fan 42 at a lowerspeed than the first spool 30. The second spool 32 includes a secondshaft 50 that interconnects a second compressor 52 and second turbine54. The first spool 30 runs at a relatively lower pressure than thesecond spool 32. It is to be understood that “low pressure” and “highpressure” or variations thereof as used herein are relative termsindicating that the high pressure is greater than the low pressure. Anannular combustor 56 is arranged between the second compressor 52 andthe second turbine 54. The first shaft 40 and the second shaft 50 areconcentric and rotate via bearing systems 38 about the engine centralaxis A which is collinear with their longitudinal axes.

The core airflow is compressed by the first compressor 44 then thesecond compressor 52, mixed and burned with fuel in the annularcombustor 56, then expanded over the second turbine 54 and first turbine46. The first turbine 46 and the second turbine 54 rotationally drive,respectively, the first spool 30 and the second spool 32 in response tothe expansion.

FIG. 2 illustrates an example airfoil 60. In this example, the airfoil60 is a turbine blade of the turbine section 28. The airfoil 60 may bemounted on a turbine disk in a known manner with a plurality of likeairfoils. Alternatively, it is to be understood that although theairfoil 60 is depicted as a turbine blade, the disclosure is not limitedto turbine blades and the concepts disclosed herein are applicable tocompressor blades in the compressor section 24, fan blades in the fansection 22 or any other blade structures. Thus, some features that areparticular to the illustrated turbine blade are to be consideredoptional.

The airfoil 60 includes an airfoil portion 62, a platform 64 and a root66. The platform 64 and the root 66 are particular to the turbine bladeand thus may differ in other airfoil structures or be excluded in otherairfoil structures.

The airfoil 60 includes a body 68 that defines a longitudinal axis Lbetween a base 70 at the platform 64 and a tip end 72. The longitudinalaxis L in this example is perpendicular to the engine central axis A.The body 68 includes a leading edge (LE) and a trailing edge (TE) and afirst side wall 74 (suction side) and a second side wall 76 (pressureside) that is spaced apart from the first side wall 74. The first sidewall 74 and the second side wall 76 join the leading edge (LE) and thetrailing edge (TE).

The airfoil portion 62 connects to the platform 64 at a fillet 80. Theplatform 64 connects to the root 66 at buttresses 82. The root 66generally includes a neck 84 and a serration portion 86 for securing theairfoil 60 in a disk.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” “circumferential,”“radial” and the like are with reference to the normal operationalattitude and engine central axis A, unless otherwise indicated.Furthermore, with reference to the engine 20, the tip end 72 of theairfoil 60 is commonly referred to as the outer diameter of the airfoil60 and the root 66 is commonly referred to as the inner diameter of theairfoil 60. The platform 64 includes an upper surface 64 a that boundsan inner diameter of a gas path, generally shown as G, over the airfoilportion 62. In this example, the tip end 72 of the airfoil body 68includes another platform 64′ that bounds the diametrically outer regionof the gas flow G.

In the illustrated example, the tip end 72 of the airfoil body 68includes at least one porous, abradable element 88. The term “abradable”as used in this disclosure refers to a structure that is less abrasivethan another, more abrasive structure which rubs against the lessabrasive structure such that the less abrasive structure will be wornaway at a greater rate than the more abrasive structure.

In this example, the airfoil body 68 includes two such porous, abradableelements 88. It is to be understood, however, that airfoil body 68 mayinclude only a single porous, abradable element 88 that extendspartially or fully across the tip end 72, or a greater number of porous,abradable elements 88. The one or more porous, abradable elements 88 arefixed on the platform 64′.

In one example, the porous, abradable element 88 has a regular,geometric structure, such as the honeycomb 90 shown in FIG. 3A. Thehoneycomb 90 includes cell walls 90 a that are configured in a hexagonalpattern, although other patterns of regular porosity could be used. Thecell walls 90 a define open, geometric pores 90 b. As shown incross-section in FIG. 3B, the cell walls 90 a of the honeycomb 90 aretapered and have a thicker base at the platform 64′ that tapers to athinner portion at the distal end. The tapering thus defines afrustoconical shape of the pores 90 b. The tapering of the cell walls 90a permits the distal ends of the cell walls 90 a to be more easily wornaway in comparison to portions of the cell walls 90 a that are closer tothe platform 64′. Additionally, the thicker base of the cell walls 90 astrengthen the honeycomb 90. Thus, the tapering can be used to tailorthe abradabilty and strength of the porous, abradable elements 88.

Alternatively, the porous, abradable element 88 has a random array ofpores 92, as shown in FIG. 3C. The random array of pores 92 may form aclosed porosity or an open, interconnected porosity.

The porous, abradable element 88 of any of the above examples has afirst composition and the airfoil body 68 has a second composition. Thefirst composition can be the same or different than the secondcomposition with respect to the chemical elements and amounts ofchemical elements present. In one example, the compositions areequivalent and are both nickel-based alloys. In another example, thecomposition of the porous, abradable element 88 is a differentnickel-alloy composition than the airfoil body 68, a different metallicalloy than the airfoil body 68 or is a non-metallic material.

As pointed out above, the airfoil 60 may be a blade within the fansection 22, the compressor section 24 or the turbine section 28.Generally, the airfoils 60 are mounted on disk in a known manner suchthat the disk includes a plurality of the airfoils 60 circumferentiallymounted around the periphery of the disk. As shown in FIG. 4, a staticshroud 96 extends circumferentially around the airfoils 60 adjacent theporous, abradable elements 88. The static shroud 96 includes at leastone edge 96 a (e.g., a “knife” edge), that extends circumferentially andradially inwardly with regard to the engine longitudinal axis A towardthe porous, abradable elements 88. The static shroud 96 may includeseveral edges 96 a that are axially offset from each other.

In operation of the engine 20, the airfoils 60 rotate around the enginecentral axis A such that the porous, abradable elements 88 contact theedge 96 a of the static shroud 96. Because the porous, abradableelements 88 are porous, the edge 96 a or edges, which are generallysolid, wear a groove in the porous, abradable elements 88. Theinteraction between the edge 96 a and the porous, abradable elements 88thus provides a dynamic seal between the moving airfoils 60 and thestatic shroud 96. Traditionally, due to manufacturing limitations in diecasting techniques used to form airfoils, such edges are provided on thetip ends of airfoils and porous elements are provided on the staticshroud. However, if edges on the airfoils 60 are circumferentiallymisaligned, the groove formed in the porous elements on the shroudbecomes enlarged and thus provides less sealing. By instead of providingthe porous, abradable elements 88 on the airfoils 60, the edge 96 a canbe made circumferentially continuous over multiple airfoils 60, as shownin FIG. 4, and controlled within a tight dimensional tolerance toprovide a tighter groove that enhances sealing. The static shroud 96 isalso easier to manufacture in comparison to shrouds that have porouselements due to elimination of assembly of a porous element on thestatic shroud 96.

FIG. 5 shows a perspective view with a portion of the static shroud 96cutaway to reveal several airfoils 60 at their tip ends 72. Theplatforms 64′ of the tip ends 72 interlock at bearing surfaces 72′ suchthat the porous, abradable elements 88 on each tip end circumferentiallyalign. The porous, abradable elements 88 are elongated along thecircumferential direction and are axially spaced apart from each otherto form sides of a channel 100 there between. The porous, abradableelements 88 on each of the platforms 64′ are also circumferentiallyoffset from each other. The platforms 64′ form a bottom of the channel,which is opposite of an open top 102 of the channel 100.

The geometries disclosed herein, such as, but not limited to, thetapered cell walls 90 a of the honeycomb 90, may be difficult to formusing conventional casting technologies. Thus, a method of processing anairfoil having the features disclosed herein includes an additivemanufacturing process, as schematically illustrated in FIG. 6. Powderedmetal suitable for aerospace airfoil applications is fed to a machine,which may provide a vacuum, for example. The machine deposits multiplelayers of powdered metal onto one another. The layers are selectivelyjoined to one another with reference to Computer-Aided Design data toform solid structures that relate to a particular cross-section of theairfoil. In one example, the powdered metal is selectively melted usinga direct metal laser sintering process or an electron-beam meltingprocess. Other layers or portions of layers corresponding to negativefeatures, such as cavities, openings or porosity, are not joined andthus remain as a powdered metal. The unjoined powder metal may later beremoved using blown air, for example. With the layers built upon oneanother and joined to one another cross-section by cross-section, anairfoil or portion thereof, such as for a repair, with any or all of theabove-described geometries, may be produced. The airfoil may bepost-processed to provide desired structural characteristics. Forexample, the airfoil may be heated to reconfigure the joined layers intoa single crystalline structure.

Additionally, the method may be used as a retrofit or a repair of theporous, abradable elements 88. In one example, the method is used toform the porous, abradable elements 88 on a prior-existing airfoil thatdoes not originally include such elements. In another example, themethod is used as a repair to rebuild worn porous, abradable elements88. For example, the worn, porous, abradable elements are machined downto a flat surface, such as the flat top surface of the platform 64′.Once machined to a flat surface, a new porous, abradable element 88 canbe built upon the flat surface using the method described above.Alternatively, a new porous, abradable element 88 can be builtseparately and then attached, such as by brazing, to the flat surface.

In a further example, the platform 64′ is provided with a sacrificiallayer 64 a′ (FIG. 2) that can be partially machined away along with aworn, porous, abradable element. The sacrificial layer 64 a′ has apredefined longitudinal thickness. As long as at least a portion of thesacrificial layer 64 a′ remains after machining away a worn, porous,abradable element, the airfoil 60 meets dimensional requirements.However, once the sacrificial layer 64 a′ is completely machined away,such as after several repairs and rebuilds of the porous, abradableelement 88, the complete consumption of the sacrificial layer 64 a′indicates that the airfoil 60 does not meet dimensional requirements andshould no longer be repaired, for example.

FIG. 7 shows a perspective view of a tip end 172 of a modified airfoil160. In this disclosure, like reference numerals designate like elementswhere appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding elements. In this example, the tip end 172 includes anopening 110 between the porous, abradable elements 88. The opening 110opens to an internal cavity within the airfoil 160. Functionally, theopening 110 allows removal of any loose powder material remaining in theinternal cavity of the airfoil after or during the manufacturing processdescribed above. The opening 110 may later be sealed over or may remainopen.

Although a combination of features is shown in the illustrated examples,not all of them need to be combined to realize the benefits of variousembodiments of this disclosure. In other words, a system designedaccording to an embodiment of this disclosure will not necessarilyinclude all of the features shown in any one of the Figures or all ofthe portions schematically shown in the Figures. Moreover, selectedfeatures of one example embodiment may be combined with selectedfeatures of other example embodiments.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. The scope of legal protection given tothis disclosure can only be determined by studying the following claims.

What is claimed is:
 1. A blade comprising: an airfoil including a baseand a free, tip end, the tip end including at least one porous,abradable element.
 2. The blade as recited in claim 1, wherein the tipend includes a platform and the at least one porous, abradable elementis on the platform.
 3. The blade as recited in claim 2, wherein the atleast one porous, abradable element includes a first porous, abradableelement and a second porous, abradable element that define sides of achannel, with the platform defining a bottom of the channel and thechannel having an open top opposite the bottom.
 4. The blade as recitedin claim 1, wherein the at least one porous, abradable element includesa honeycomb.
 5. The blade as recited in claim 4, wherein the honeycombhas cell walls that are tapered.
 6. The blade as recited in claim 1,wherein the at least one porous, abradable element includes a randomarray of pores.
 7. The blade as recited in claim 1, wherein the at leastone porous, abradable element includes geometric pores.
 8. The blade asrecited in claim 1, wherein the at least one porous, abradable elementincludes a first porous, abradable element and a second porous,abradable element spaced apart from the first porous, abradable element.9. The blade as recited in claim 8, wherein the first porous, abradableelement and the second porous, abradable element are elongated along thesame direction and offset from each other along the direction.
 10. Theblade as recited in claim 1, wherein the at least one porous, abradableelement has a first composition and the airfoil has a second compositionthat is equivalent to the first composition.
 11. The blade as recited inclaim 1, wherein the at least one porous, abradable element has a firstcomposition and the airfoil has a second composition that is differentthan the first composition.
 12. A turbine engine comprising: optionally,a fan; a compressor section; a combustor in fluid communication with thecompressor section; and a turbine section in fluid communication withthe combustor, the turbine section being coupled to drive the compressorsection and the fan, and at least one of the fan, the compressor sectionand the turbine section including a plurality of blades, each of theplurality of blades having an airfoil including a base and a free, tipend, the tip end including at least one porous, abradable element. 13.The turbine engine as recited in claim 12, further including a staticshroud extending circumferentially around the plurality of bladesadjacent the porous, abradable elements, and the static shroud includesat least one edge extending circumferentially and radially inwardlytoward the porous, abradable elements.
 14. The turbine engine as recitedin claim 13, wherein the at least one edge is circumferentiallycontinuous over multiple ones of the plurality of blades.
 15. A methodfor processing a blade, the method comprising: depositing multiplelayers of a powdered metal onto one another; joining the layers to oneanother with reference to data relating to a particular cross-section ofa blade; and producing the blade with an airfoil including a base and afree, tip end, the tip end including at least one porous, abradableelement.
 16. A method for processing a blade, the method comprising:providing a blade having an airfoil including a base and a free, tipend; and forming at the tip end at least one porous, abradable element.17. The method as recited in claim 16, further including, prior to theforming, removing a prior porous, abradable element from the tip end.18. The method as recited in claim 17, wherein the removing includesmachining the tip end to form a flat surface.
 19. The method as recitedin claim 18, wherein the forming includes forming the at least oneporous, abradable element on the flat surface.
 20. The method as recitedin claim 16, wherein the forming includes depositing multiple layers ofa powdered metal on to one another and joining the layers to one anotherwith reference to data relating to a particular cross-section of the atleast one porous, abradable element.